FIG. 1 shows a ducted fan gas turbine engine 10 comprising, in axial flow series: an air intake 12, a propulsive fan 14 having a plurality of fan blades 16, an intermediate pressure compressor 18, a high-pressure compressor 20, a combustor 22, a high-pressure turbine 24, an intermediate pressure turbine 26, a low-pressure turbine 28 and a core exhaust nozzle 30. The fan, compressors and turbine are all rotatable about a principal axis 31 of the engine 10. A nacelle 32 generally surrounds the engine 10 and defines the intake 12, a bypass duct 34 and a bypass exhaust nozzle 36.
Air entering the intake 12 is accelerated by the fan 14 to produce a bypass flow and a core flow. The bypass flow travels down the bypass duct 34 and exits the bypass exhaust nozzle 36 to provide the majority of the propulsive thrust produced by the engine 10. The core flow enters in axial flow series the intermediate pressure compressor 18, high pressure compressor 20 and the combustor 22, where fuel is added to the compressed air and the mixture burnt. The hot combustion products expand through and drive the high, intermediate and low-pressure turbines 24, 26, 28 before being exhausted through the nozzle 30 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines 24, 26, 28 respectively drive the high and intermediate pressure compressors 20, 18 and the fan 14 by interconnecting shafts 38, 40, 42.
The performance of gas turbine engines, whether measured in terms of efficiency or specific output, is generally improved by increasing the turbine gas temperature. It is therefore desirable to operate the turbines at the highest possible temperatures. As a result, the turbines in state of the art engines, particularly high pressure turbines, operate at temperatures which are greater than the melting point of the material of the blades and vanes making some form of cooling necessary. However, increasing cooling of components generally represents a reduction in turbine efficiency and so much effort is spent in finding a satisfactory trade-off between turbine entry temperature, the life of a cooled turbine component and specific fuel consumption. This has led to a great deal of research and development of new materials and designs which can allow an efficient increase of the gas turbine entry temperature.
The present invention seeks to provide improved cooling arrangements for a gas turbine.